This invention relates generally to gas turbine engines and more particularly to slot cooled combustor liners used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Combustors used in aircraft engines typically include inner and outer combustor liners. The liners contain the combustion process and facilitate the distribution of air to the various combustor zones in prescribed amounts.
Because they are exposed to intense heat generated by the combustion process, combustor liners are cooled to meet life expectancy requirements. Liner cooling is commonly provided by diverting a portion of the compressed air (which is relatively cool) and causing it to flow over the outer surfaces of the liners. In addition, a thin layer of cooling air is provided along the combustion side of the liners by directing cooling air flow through cooling holes formed in the liners. This technique, referred to as film cooling, reduces the overall thermal load on the liners because the mass flow through the cooling holes dilutes the hot combustion gas next to the liner surfaces, and the flow through the holes provides convective cooling of the liner walls.
In one known configuration, film cooled combustor liners include a series of connected panel sections with a bump or nugget formed on the forward end of each panel section. Each nugget has an axially oriented slot formed on the hot gas side thereof, and a plurality of cooling holes is formed in each nugget. Compressor discharge air passes through the cooling holes and exits the cooling slots to produce the film of cooling air on the hot gas side of the corresponding panel section.
It is also known to use a thermal barrier coating on the hot gas side of combustor liners for providing further protection from the high temperatures of combustion. Thermal barrier coatings typically comprise a bond coat having a composition similar to the base material and a ceramic top coat.
Each cooling slot defines an overhanging portion, also known as a cooling lip, which is exposed to the hot combustion gas. Cooling lips must be sufficiently thick for structural reasons. As a result, cooling lips have bluff trailing edge surfaces of significant thickness relative to cooling slot height. Because of the inherently separated flow off this bluff trailing edge surface, and the fact that the surface has a significant view factor to the combustion flame, there is substantial heat input to the trailing edge surface. The gas conditions on the trailing edge surface typically comprise a mix between the cooling air temperature and the hot gas temperature, thereby further adding to the heat input. Thus, despite the fact that jets of air from the cooling holes cool the cooling lip generally, the heat input through the bluff trailing edge surface can significantly increase liner metal temperatures and reduce service life.
To reduce the effects of trailing edge heat input, it is known to taper the cooling lip such that the thickest portion is at the base and the thinnest portion is at the trailing edge. This configuration reduces the trailing edge surface thickness while maintaining structural stability to resist buckling. Taper angles between 5 and 10 degrees from parallel are commonly used. With such taper angles, conventional liners have trailing edge thicknesses of approximately 0.050 to 0.060 inches. Thus, while tapering the cooling lip reduces the thickness of the bluff trailing edge surface, the surface is still relatively thick. Furthermore, increasingly thicker ceramic top coats are being used in thermal barrier coatings as a way of extending the life of combustor liners. This has the effect of increasing the overall thickness of the bluff trailing edge surface, thereby increasing the exposed area of the trailing edge. It also increases mainstream turbulence at the cooling slot exit, which degrades the downstream effectiveness of the cooling film.
It would be theoretically possible to achieve thinner trailing edge thickness by applying a sharper than typical taper angle to the cooling lip. However, this would result in a significant portion of the lip being structurally weakened and a significant length of the lip being too thin to manufacture practically without deforming the material. The weakened structure would be susceptible to thermal buckling under hot streaking conditions.
Accordingly, it is desirable to reduce the thickness of the cooling lip trailing edge bluff surface without diminishing the structural integrity of the cooling lip.
The above-mentioned need is met by the present invention which provides a gas turbine combustor liner that includes an annular panel section having a cooling nugget formed on one end thereof. An annular cooling lip is formed on the cooling nugget so as to define a cooling slot. The cooling lip includes a hot side, a cold side and a trailing edge surface. The hot side defines a compound taper for minimizing the thickness of the trailing edge surface. In one embodiment, the compound taper includes a first tapered surface disposed on a forward portion of the cooling lip and a second tapered surface disposed on an aft portion of the cooling lip, wherein the second tapered surface has a greater taper than the first tapered surface.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.